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| United States Patent |
6,209,312
|
|
Singer
, et al.
|
April 3, 2001
|
Rocket motor nozzle assemblies with erosion-resistant liners
Abstract
A nozzle assembly including a nozzle structure and liner is disclosed. The
nozzle structure is made of at least one carbon-based material and
includes a nose tip region, a restricted cross-sectional throat region,
and an exit cone region that collectively provide an interior surface
configured to define a converging-diverging pathway. The liner includes
leg and body portions. The leg portion protrudes over an edge or into a
groove of the nozzle structure to engage the liner to the nozzle
structure. The body portion of the liner covers at least the throat region
of the nozzle structure along the flow path to obstruct high temperature
combustion products from causing recession of the nozzle structure. The
erosion-resistant liner has at least one irregularity that extends, in a
continuous manner, radially at least along the leg portion and,
optionally, longitudinally along the body portion. The irregularity is
constructed and arranged to permit thermal deformation of the liner in
response to thermally induced hoop stresses encountered in the nozzle
structure during motor operation so as to reduce or eliminate thermal
fracturing of the liners. Suitable irregularities include, by way of
example, clearances in a circumferentially discontinuous liner.
| Inventors:
|
Singer; Victor (Newark, DE);
Carr, Jr.; Clyde E. (New London, PA)
|
| Assignee:
|
Cordant Technologies Inc (Salt Lake City, UT)
|
| Appl. No.:
|
288330 |
| Filed:
|
April 8, 1999 |
| Current U.S. Class: |
60/770; 239/265.11 |
| Intern'l Class: |
F02K 001/00 |
| Field of Search: |
60/271
239/265.11,397.5,591
|
References Cited
U.S. Patent Documents
| 2987874 | Jun., 1961 | Nicholson | 60/35.
|
| 3048972 | Aug., 1962 | Barlow | 60/271.
|
| 3137995 | Jun., 1964 | Othmer et al. | 60/35.
|
| 3165864 | Jan., 1965 | Shulze | 50/464.
|
| 3189477 | Jun., 1965 | Shaffer | 117/46.
|
| 3194013 | Jul., 1965 | Dagneau et al. | 60/35.
|
| 3200585 | Aug., 1965 | Climent et al. | 60/35.
|
| 3226929 | Jan., 1966 | McKenna | 60/35.
|
| 3228186 | Jan., 1966 | Allen | 60/35.
|
| 3261558 | Jul., 1966 | Davies | 239/601.
|
| 3545679 | Dec., 1970 | McAllister | 239/265.
|
| 3723214 | Mar., 1973 | Meraz, Jr. | 156/87.
|
| 3762644 | Oct., 1973 | Mikeska | 239/127.
|
| 4477024 | Oct., 1984 | O'Driscoll et al. | 239/265.
|
| 4729512 | Mar., 1988 | Laing | 239/265.
|
| 4917968 | Apr., 1990 | Tuffias et al. | 428/621.
|
| 5263349 | Nov., 1993 | Felix et al. | 72/38.
|
| 5802842 | Sep., 1998 | Hook et al. | 60/271.
|
Other References
Singer et al., Benchmark Issues in Design of Metal High-Performance
Pressure Vessels, Mar. 1998, p. 12.
Singer et al., "Application of Fracture Mechanics In Design and Analysis Of
Pressure Vessels", Engineering Fracture Mechanics, 1969, vol. 1, pp.
507-517.
|
Primary Examiner: Freay; Charles G.
Attorney, Agent or Firm: Sullivan Law Group
Goverment Interests
GOVERNMENT LICENSE RIGHTS
The U.S. Government has a paid-up license in this invention and the right
in limited circumstances to require the patent owner to license others on
reasonable terms as provided for by the terms of contract no
N60921-95-C0032 awarded by the Department of the Navy.
Parent Case Text
RELATED APPLICATION
Priority is claimed based on provisional application 60/081,184 filed in
the U.S. Patent & Trademark Office on Apr. 9, 1998, the complete
disclosure of which is incorporated herein by reference.
Claims
We claim:
1. A nozzle assembly mountable to a rocket motor body to form part of a
rocket motor assembly and constructed and arranged so that, during
operation of the rocket motor assembly, said nozzle assembly receives high
temperature combustion products from a combustion chamber of the rocket
motor body and discharges the high temperature combustion products to
propel and/or divert the rocket motor assembly, said nozzle assembly
comprising:
a mount structure constructed and arranged to permit mounting of said
nozzle assembly to the rocket motor body;
a nozzle structure associated with said mount structure and comprised of at
least one carbon-based material, said nozzle structure comprising a nose
tip region, a restricted cross-sectional throat region, and a exit cone
region that collectively provide an interior surface configured to define
a converging-diverging flow path through which the combustion products
pass during operation of the rocket motor assembly with one or more edges
and/or grooves formed in said interior surface; and
one or more erosion-resistant liners respectively comprising at least one
leg portion and a corresponding body portion angled relative to said
corresponding leg portion, each of said leg portions protruding into
respective ones of said edges or grooves of said nozzle structure to
engage said liners to said nozzle structure, said body portion or body
portions of said liners collectively covering at least said throat region
so that said body portions obstruct the high temperature combustion
products from coming into sufficient contact with said nozzle structure to
cause said nozzle structure to recede during operation of the rocket motor
assembly,
wherein each of said liners have one or more irregularities extending
radially along said leg portion, said irregularities constructed and
arranged to permit thermal deformation of said liners in response to
thermally induced hoop stresses encountered in said nozzle structure
during motor operation so as to reduce or eliminate thermal fracturing of
said liners.
2. The nozzle assembly of claim 1, wherein said liners comprise at least
one tungsten alloy.
3. The nozzle assembly of claim 2, wherein said tungsten alloy is selected
from the group consisting of tungsten rhenium, tungsten hafnium, and
tungsten tantalum.
4. The nozzle assembly of claim 1, wherein said liners comprise at least
one carbide of a refractory metal.
5. The nozzle assembly of claim 4, wherein said carbide is selected from
the group consisting of hafnium carbide, tantalum carbide, and zirconium
carbide.
6. The nozzle assembly of claim 1, wherein said liners comprise at least
one boride of a refractory metal.
7. The nozzle assembly of claim 6, wherein said boride is selected from the
group consisting of hafnium diboride and zirconium diboride.
8. The nozzle assembly of claim 1, wherein at least one of said liners
comprises a plurality of stacked layers, an outermost one of said stacked
layers being chemically inert with respect to said nozzle structure.
9. The nozzle assembly of claim 8, wherein one of said stacked layers that
is radially inside of said outermost layer comprises tungsten.
10. The nozzle assembly of claim 1, wherein said irregularities are defined
by circumferential discontinuities.
11. The nozzle assembly of claim 10, wherein said liners have one or more
slits therein, each of said slits extending radially over a portion of
said liners.
12. The nozzle assembly of claim 1, wherein said liners respectively
comprise a plurality of circumferentially discontinuous segments
collectively forming a ring with clearances defined between adjacent ones
of said circumferentially discontinuous segments.
13. The nozzle assembly of claim 12, wherein said liners respectively
comprise a plurality of stacked layers, an outermost one of said stacked
layers being chemically inert with respect to said nozzle structure.
14. The nozzle assembly of claim 13, wherein said clearances of adjacent
ones of said stacked layers are circumferentially staggered relative to
each other.
15. The nozzle assembly of claim 14, further comprising at least one
adhesive disposed in said clearances between said segments.
16. A nozzle assembly mountable to a rocket motor body to form part of a
rocket motor assembly and constructed and arranged so that, during
operation of the rocket motor assembly, said nozzle assembly receives high
temperature combustion products from a combustion chamber of the rocket
motor body and discharges the high temperature combustion products to
propel and/or divert the rocket motor assembly, said nozzle assembly
comprising:
a mount structure constructed and arranged to permit mounting of said
nozzle assembly to the rocket motor body;
a nozzle structure associated with said mount structure and comprised of at
least one carbon-based material, said nozzle structure comprising a nose
tip region, a restricted cross-sectional throat region, and an exit cone
region that collectively provide an interior surface configured to define
a converging-diverging flow path through which the combustion products
pass during operation of the rocket motor assembly with one or more edges
and/or grooves formed in said interior surface; and
one or more erosion-resistant liners respectively comprising at least one
leg portion and a corresponding body portion angled relative to said
corresponding leg portion, each of said leg portions protruding into
respective ones of said edges or grooves of said nozzle structure to
engage said liners to said nozzle structure, said body portion or body
portions of said liners collectively covering said throat region and
optionally said nose tip region along said flow path and said exit cone
region along a section of said flow path that is prone to more than
negligible amount of recession so that said body portions obstruct the
high temperature combustion products from coming into sufficient contact
with said nozzle structure to cause said nozzle structure to recede during
operation of the rocket motor assembly,
wherein each of said liners have one or more irregularities in the form of
grooves extending radially along said leg portion and optionally
longitudinally along said body portion, said irregularities constructed
and arranged to permit thermal deformation of said liners in response to
thermally induced hoop stresses encountered in said nozzle structure
during motor operation so as to reduce or eliminate thermal fracturing of
said liners.
17. The nozzle assembly of claim 16, wherein said liners comprise at least
one tungsten alloy.
18. The nozzle assembly of claim 17, wherein said tungsten alloy is
selected from the group consisting of tungsten rhenium, tungsten hafnium,
and tungsten tantalum.
19. The nozzle assembly of claim 16, wherein said liners comprise at least
one carbide of a refractory metal.
20. The nozzle assembly of claim 19, wherein said carbide is selected from
the group consisting of hafnium carbide, tantalum carbide, and zirconium
carbide.
21. The nozzle assembly of claim 16, wherein said liners comprise at least
one boride of a refractory metal.
22. The nozzle assembly of claim 21, wherein said boride is selected from
the group consisting of hafnium diboride and zirconium diboride.
23. The nozzle assembly of claim 16, wherein grooves extend longitudinally
along said body portion.
24. The nozzle assembly of claim 16, wherein said liner covers said nose
tip region.
Description
BACKGROUND OF THE INVENTION
1. Field of the Invention
This invention relates to a rocket motor nozzle assembly designed to reduce
nozzle recession, especially for rocket motor nozzle assemblies comprising
carbon-based materials.
2. Description of the Related Art
One factor taken into consideration in nozzle design is the configuration
of the divergent/convergent pathway defined by the nozzle, and the
restrictive orifice or throat area of the nozzle. Mass flow through the
nozzle produces the force or thrust produced by the rocket motor. The
proportions of the mass flow pathway, particularly the ratio of area at
the exit plane to area at the throat, establish how efficiently the nozzle
converts pressure in the mass flow stream to thrust produced by the motor.
It is within the purview of those skilled in the art to design a nozzle to
optimize the ratio of exit area to throat area.
Another factor taken into consideration in nozzle design is the weight
penalty imparted to the rocket motor by the nozzle assembly. Intuitively,
a lesser weight nozzle assembly is desirable because the lesser the weight
of the nozzle assembly, the farther the rocket motor assembly can travel.
For this reason, carbon-based materials are highly advantageous for use as
nozzle insulation due to their low weight. As referred to herein,
carbon-based materials include, but are not limited to, carbon or graphite
bulk and composite materials with constituents previously subject to
carbonization or graphitization, known as carbon/carbon, graphite/carbon,
and cloth, fiber, or powder-filled phenolic composites, and also a large
array of metal or silicon carbides.
It is widely acknowledged in the industry, however, that carbon-based
nozzle throats tend to recede, especially at high operating temperatures
and pressures. Studies have identified several reasons for nozzle throat
recession. One among several reasons involves oxidation reactions between
the carbon-based nozzle material and oxygen-containing constituents of the
combustion products. As the propellant in the rocket motor burns, the
carbon-based nozzle is exposed to the hot combustion gases, such as
O.sub.2, H.sub.2 O, CO.sub.2, and NO. These gases, especially water and
carbon dioxide, tend to react with the carbon-based nozzle materials to
produce carbon monoxide gas, which is carried off with the discharged
combustion products. Another reason for nozzle throat erosion is the
abrasive impact of high velocity particles in the gas stream against the
nozzle throat. Still another reason for nozzle throat recession involves
pyrolysis of the throat material itself, particularly when volatile
decomposition products form.
The recession of the nozzle throat inner surface during motor operation is
a source of several problems in rocket operation. As the nozzle throat
material recedes, the exit area to throat area ratio (or expansion ratio)
diminishes, thereby decreasing the efficiency of the nozzle. Additionally,
rough nozzle surfaces, which tend to form during nozzle recession, have
been shown to undergo recession at faster rates than smooth surfaces.
Thus, the nozzle throat recession process can be characterized as a
self-perpetuating phenomenon. Another problem attributable to nozzle
recession is a loss of predictability. Calculations for determining
acceptable payloads and requisite propellant grain stocks must be accurate
to ensure that the rocket will reach its intended target. The calculations
necessary for ascertaining rocket dimensions and payloads are dependent
upon many variables, including nozzle throat dimension. In-flight
variations of nozzle throat dimension due to recession can significantly
complicate, if not render impossible, precise motor performance
calculations.
To address the shortcomings of carbon-based nozzle throats, refractory
metal and metal alloys are occasionally used in spite of their high
specific gravities. Examples of such materials are tungsten and its
alloys.
However, the weight penalty and expense associated with the presence of the
tungsten and other refractory metals often make these refractory materials
impractical and uneconomical for applications involving bulky throat
insert cross-section sizes. Additionally, such cross sections are subject
to tensile and compressive stresses due to thermal shock early in motor
burn when thermal expansions near the rapidly heated exposed surfaces are
restrained by cooler regions of the cross section farther from the exposed
surfaces. Indeed, surface heating can be so intense that temperature
gradients of thousands of degrees per inch are possible. Such thermal
stresses in both the axial and tangential (or hoop) directions can produce
thermal fractures in the nozzle component, and potentially ejection from
the motor.
Refractory ceramic materials have also occasionally been used in an attempt
to address the shortcomings of carbon-based nozzles. Refractory ceramics
present lesser weight penalties than refractory metals, but thermal shock
penalties may be greater than in refractory metals.
To address the problem of restrained thermal deformations, U.S. Pat. No.
3,200,585 to Climent et al., the complete disclosure of which is
incorporated herein by reference, discloses the use of a plurality of
tungsten washers as constituting the throat portion of a nozzle, with the
washers being stacked to form a cylindrical structure. The washers include
radially extending slits to permit expansion and contraction of the
washers in response to thermal stresses. However, the tubular structure
disclosed by Climent et al. has several drawbacks. For example, the
tubular structure is disposed only in the throat region of the nozzle
pathway, leaving other portions of the pathway, such as the nosetip,
entrance, and susceptible exit regions of the pathway, unprotected.
Additionally, tungsten is chemically reactive with carbon-based insulation
or substrate. As a consequence, the tungsten and carbon-based portions are
prone to recession. Finally, the mounting of the tungsten washers and
their support member can be a laborious and time-consuming process.
It would, therefore, be a significant advancement in the art to provide a
nozzle assembly that takes advantage of the low weight of carbon-based
materials and the erosion resistance of metals and alloys, yet does not
impose an undue weight penalty to the rocket motor assembly and avoids
nozzle recession and its associated problems.
SUMMARY OF THE INVENTION
It is, therefore, an object of this invention to provide a rocket motor
assembly that accomplishes the above-mentioned improvement in the art.
In accordance with the principles of this invention, these and other
objects are attained by the provision of a nozzle assembly mountable to a
rocket motor body to form part of a rocket motor assembly for receiving
and discharging high temperature combustion products from a combustion
chamber of the rocket motor body. In accordance with an embodiment of this
invention, the nozzle assembly includes at least a mount structure adapted
for mounting of the nozzle assembly to the rocket motor body, a
converging-diverging nozzle structure associated with the mount structure
and comprised of at least one carbon-based material, and one or more
erosion-resistant liners. The nozzle structure comprises a nose tip
region, a restricted cross-sectional throat region, and an exit cone
region that collectively provide an interior surface configured to define
a converging-diverging flow path through which the combustion products
pass during operation of the rocket motor assembly. Each of the liners
comprises at least one leg portion and a corresponding body portion angled
relative to the leg portion. Each of the leg portions of the liners
protrudes into an edge or groove of the nozzle structure to engage the
liner to the nozzle structure. The body portions of the liners
collectively cover the throat region and optionally the nose tip region of
the nozzle structure along the flow path, as well as the exit cone region
along sections of the flow path that are prone to more than negligible
amounts of recession. In this manner, the body portions of the liners
obstruct high temperature combustion products from coming into sufficient
contact with the nozzle structure to cause the underlying nozzle structure
to recede during operation of the rocket motor assembly.
The erosion-resistant liners each have at least one irregularity that
extends, in a continuous manner, radially along the leg portion and,
optionally, longitudinally along the body portion of the liner. The
irregularity is or irregularities are constructed and arranged to permit
rather than restrain thermal deformation of the liners in response to
non-uniform heating of the nozzle structure during motor operation so as
to reduce (compared to had the irregularities not been present) or
eliminate fracturing of the liners which can occur when the deformations
are restrained. Suitable irregularities include, by way of example,
clearances formed from a circumferentially discontinuous liner, and/or
protrusions or corrugations extending across the liners.
The inventive nozzle assemblies may be applied to various kinds and types
of rocket motors having one-time burn or multiple-burn duty cycles.
Representative rocket motors in which the inventive nozzle assemblies can
be applied include, by way of example, satellite propulsion motors,
tactical motors, and divert/attitude control motors.
These and other objects, features, and advantages of the present invention
will become apparent from the following detailed description of the
invention when taken in conjunction with the accompanying drawings which
illustrate, by way of example, the principles of this invention.
BRIEF DESCRIPTION OF THE DRAWINGS
The accompanying drawings are presented for the purpose of elucidating the
principles of this invention. In the drawings:
FIG. 1 is a schematic depicting in cross-sectional view a portion of a
rocket nozzle having liners with radial irregularities formed therein;
FIG. 2 is a schematic depicting in isolated cross-sectional side view a
two-component liner in accordance with a first embodiment of the present
invention, in which the irregularities are clearances;
FIG. 3 is a schematic depicting in plan view a component of the liner
illustrated in FIG. 2;
FIG. 4 is a perspective view of a liner in accordance with a second
embodiment of this invention, in which the irregularities are grooves;
FIG. 5 is a schematic of a plan view of the liner of FIG. 4;
FIG. 6 is a sectional view of the a portion of the liner of FIGS. 4 and 5;
and
FIG. 7 is a schematic depicting in cross-sectional view a rocket motor
having a conventional rocket nozzle to which the liner of the present
invention is capable of being applied.
DETAILED DESCRIPTION OF THE INVENTION
Referring now more particularly to the drawings, there is shown in FIG. 7 a
conventional rocket motor assembly, generally designated by reference
numeral 70, which a skilled artisan having reference to this disclosure
could modify to include the inventive nozzle liner. The rocket motor
assembly 70 includes a case 71 enclosing a combustion chamber (unnumbered)
substantially filled by center-perforated solid propellant grain 72. The
case 71 and the solid propellant grain 72 are separated by a thin
insulating layer (unnumbered). Centrally located within the forward end of
the center perforation 73 of the propellant grain 72 of the rocket motor
assembly 70 is an igniter assembly 74 for initiating burn of the solid
propellant grain 72 along the center perforation 73.
The center perforation 73 of the solid propellant grain 72 is in fluid
communication with a nozzle assembly, which is generally designated by
reference numeral 75. The nozzle assembly 75 is mounted to the case 71 by
a mount structure 76, which in the illustrated embodiment includes, among
other things, a closure module, an aft polar boss, and insulation. The
nozzle assembly 75 includes a throat region 77, a nose tip region 78
arranged forward of the throat region 77, and an exit cone region 79
extending aftward from the throat region 77. The nose tip region 78,
throat region 77, and exit cone region 79 collective define a
converging-diverging flow path or pathway.
In operation, the solid propellant grain 72 is ignited by the igniter
assembly 74, producing combustion products that flow from the combustion
chamber into the nozzle assembly 75 and through the nose tip region 78 and
the throat region 77, before being expelled through the diverging exit
cone region 79.
The inventive nozzle assembly is shown in FIG. 1 and generally designated
by reference numeral 10. The nozzle assembly includes a carbon-based shell
structure comprised of forward annular region 11 and aft annular region
12. Representative materials for making the carbon-based structure
include, by way of example, carbon/carbon composites, graphite/carbon
composites, phenolics, and other known materials that are prone to nozzle
recession.
Extending over the pathway defining surface of region 11 is a body portion
14 of forward liner 13. The body portion 14 of the liner 13 covers the
nose tip region and part of the throat region along he flow path. The aft
annular region 12 has its pathway-defining surface covered by a body
portion 17 of an aft liner 16. The body portion 17 covers part of the
throat region, overlapping with body region 14. The body portion 17 also
covers a section of the exit cone region that is prone to more than a
negligible amount of recession. As referred to herein, a negligible amount
of recession means recession that is sufficiently small as to not affect
the impulse of the rocket motor (the product of thrust and time over the
burn duration) significantly mean. In this sense, a significant effect is
one that is a greater burden than the cost of eliminating the burden, if
the burden can be eliminated. By collectively extending over the shell
structure along the flow path, the body portions 14 and 17 of liners 13
and 16 obstruct high temperature combustion products from coming into
sufficient contact with the carbon-based material of the shell structure
and causing the recession of the shell structure during rocket motor
assembly operation.
The forward liner 13 also includes a leg portion 15 received at an edge of
the forward annular region 11. Likewise, the aft liner 16 has a leg
portion 18 received in a groove formed between the forward and aft annular
regions 11 and 12. The placement of leg portions 15 and 18 in regions 11
and 12 serves to retain the liners 13 and 16 against the nozzle structure.
Since no bonding agent or brazing technique is required between liners 13
and 16, the liners 13 and 16 are permitted to deform during motor
operation to diminish or even eliminate the adverse effects of thermal
stresses. Also, the connection provided by the leg portions 15 and 18
permit pyrolysis gases from regions 11 and 12 to escape, thereby
preventing unwanted reactions and pressure build ups. It is understood
that although bonding agents, brazing techniques, and the like are not use
in a preferred embodiment, these agents and techniques are also not
precluded from the scope of the invention. Additionally, known mechanical
fasteners and techniques can be used.
Referring now to FIGS. 2 and 3, there is shown a two-layer (or
two-component) liner 20 in isolation in accordance with a first embodiment
of this invention. The illustrated liner 20 comprises an outer layer or
component 21 with an inner layer or component 24 nestled therein. The
outer component 21 includes a substantially cylindrical body portion 22
and an annular flat leg portion 23, which is integrally formed with the
body portion 22. Similarly, the inner component 24 includes a
substantially cylindrical body portion 25 and an annular flat leg portion
26, which is integrally formed with the body portion 25. The outer
diameter of the substantially cylindrical body portion 22 should be
approximately the same, but slightly smaller than, the inner diameter of
the substantially cylindrical body portion 25 to permit the outer
component 21 to be nestled within the inner component 21.
FIG. 3 represents a plan view of the inner component 24 taken along line
III--III in FIG. 2. As shown in FIG. 3, the inner component 24 is formed
from four identical 90.degree. segments 28, 30, 32, and 34, which are
spaced from each other by irregularities in the form of clearances 36 to
thereby collectively provide the component 24 with the configuration of a
circumferentially discontinuous ring. Each of the component segments 28,
30, 32, and 34 has two radially extending slits 38 formed therein. The
clearances 36 and slits 38 are arranged at 30.degree. intervals. More
specifically, as shown in FIG. 3, the clearances 36 are arranged at the
0.degree., 90.degree., 180.degree., and 270.degree. positions, whereas the
slits 38 are arranged at the 30.degree., 60.degree., 120.degree.,
150.degree., 210.degree., 240.degree., 300.degree., and 330.degree.
positions. Similarly, the outer component 21 is preferably formed from
four identical 90.degree. segments having slits at 30.degree. intervals,
with the clearances and slits of the outer component 21 preferably being
offset by 15.degree. to the clearances and slits of the inner component
24. The offsetting of the clearances and slits of adjacent components 21
and 24 prevents gases from contacting the covered regions of the nozzle
structure, thereby eliminating local erosion.
At least during the early stages of rocket motor operation, the
substantially cylindrical body portions 22 and 25 will be hotter than the
flat leg portions 23 and 26, and, therefore, will experience greater
thermal expansion. The presence of radial clearances 36 and slits 38
reduces thermal stresses, especially hoop stresses, encountered in the
liner components 21 and 24 during motor operation by permitting, rather
than constraining, thermal expansion of the components 21 and 24.
A liner in accordance with a second embodiment of the invention is
illustrated in FIGS. 4-6. In the embodiment illustrated in FIGS. 4-6, the
circumferentially continuous liner 40 has a substantially cylindrical body
portion 42 and leg portion 44, which are angled relative to each other and
integrally formed with each other. In contrast to the clearances shown in
FIG. 3, the body and leg portions 42 and 44 of the liner 40 contain, as
longitudinally-extending and radially-extending continuous irregularities,
respectively, four channels (or creases) 46 spaced from each other by
90.degree.. Although rounded channels are shown in the illustrated
embodiment, especially in FIG. 6, it is understood that the channels could
have planar sides, and that more than one channel could be arranged next
to each other to define corrugations or other similar structures that are
thermally deformable in response to hoop stresses.
As the liner 40 is subject to thermal stresses during rocket motor
operation, the channels 46 permit the body and leg portions 42 and 44 of
the liner 40 to thermally deform, thereby reducing or eliminating thermal
fracturing of the liners. This embodiment is characterized by several
advantages, including that a second layer may be omitted, since the
presence of channels 46 does not interrupt the continuous structure of the
liner. Also, the presence of only one layer, instead of a multi-component
liner, is efficient from cost, weight, and volume considerations.
The thickness of the material should be smaller than the critical thickness
at operation temperature so as to permit thermal deformation in response
to hoop stresses. In this sense, "critical thickness" is infinitesimally
larger than what would permit the material to be folded without fracture.
For example, tungsten liners could have thicknesses as high as about 0.040
inch (about 0.10 cm).
Deep-draw and flow-form techniques, well known among those skilled in the
art of metal forming, can be used for the components shown in FIGS. 2 and
3. Plasma spray deposition techniques for making the liner 40 are well
known in the art. Machining (a group of metal removal techniques) is also
well known for those skilled in the art. For these reasons, none of the
above-mentioned techniques are described herein.
The liners are preferably made of materials resistant to erosion up to
operating temperatures of at least 5000.degree. F. (2760.degree. C.), more
preferably at least 5500.degree. F. (3038.degree. C.), still more
preferably at least 6100.degree. F. (3371.degree. C.). The liner materials
themselves need not have such temperature capabilities; instead, lesser
temperature capabilities are permissible when mitigating conditions are
present, such as, for example, when the supporting structure is an
effective heat sink or when burn rate duration is short.
Representative materials from which the liners can be made include the
following: tungsten; tungsten alloys, such as tungsten hafnium, tungsten
tantalum, and tungsten rhenium; and refractory ceramics and
fiber-reinforced ceramic composites using refractory metal compounds, such
as carbides or nitrides of hafnium, tantalum, and zirconium, and borides
of hafnium and zirconium. The inner and outer components of a liner can be
made of the same or different materials. However, where tungsten is used
as a material for forming a component of the liner, the liner is
preferably of a multi-layered construction, with a component that is not
chemically reactive with the carbon-based nozzle structure or the tungsten
being interposed between the tungsten component and the nozzle structure
to prevent reaction therebetween. Representative non-reactive component
materials that can be used in combination with tungsten components include
oxides or carbides of hafnium, tantalum, tungsten, rhenium, titanium, and
zirconium.
In order to reduce or eliminate reaction between the tungsten and
carbon-based region of the nozzle structure, the interior surface of the
nozzle structure can be sprayed with rhenium, or the contacting surface of
the tungsten layer can be oxidized.
The interface between the carbon-based shell structure and liners and the
interface between components of liners can be filled with an adhesive.
Representative adhesives include an epoxy adhesive composition, such as EA
934 available from Hysol.
As mentioned above, the inventive nozzle assemblies can be applied in
various rocket motors, including, by way of example, in satellite
propulsion motors, tactical motors, and divert/attitude control motors. In
accordance with one contemplated use of the invention, the components of
the liners have the following dimensions. The cylindrical body portions of
the liner components can have a thickness in a range of from 10 mils to 60
mils (0.01 inches to 0.06 inches, or about 0.025 cm to about 0.15 cm). The
thickness of the flat leg portions generally should not be smaller than
that of the cylindrical body portion, but may be thicker than the body
portions, particularly if the thickness of the body portion is towards the
lower end of the above-mentioned range.
Various modifications and equivalent arrangements to those illustrated fall
within the scope of the invention. For example, despite the depiction of
two liners 13 and 16 in FIG. 1, it is understood that the entire pathway
can be covered by only one liner, or by more than two liners.
Although in the embodiment illustrated in FIGS. 2 and 3 the liner 20 is
formed from two nestled components 22 and 24, it is understood that the
liner can be formed from a single component or more than two components.
Additionally, while the illustrated embodiment includes four segments,
each subtending an arc of about 90.degree., it is understood that more or
fewer segments can be used, and that the segments may subtend different
angles from each other. For example, the liners can be made of a single
segment having ends opposing each other across a radial clearance, with
radial slits optionally provided in the segment.
The foregoing detailed description of the embodiments of the invention has
been provided for the purposes of illustration and description. It is not
intended to be exhaustive or to limit the invention to the precise
embodiments disclosed. The embodiments were chosen and described to best
explain the principles of the invention and its practical application,
thereby enabling others skilled in the art to understand the invention.
While the invention has been described in connection with what is
presently considered to be the most practical and preferred embodiments,
it is understood that the invention is not limited to the disclosed
embodiments. To the contrary, this invention is intended to cover various
modifications and equivalent arrangements included within the spirit and
scope of the appended claims.
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